This invention relates to a passive clearance control system in a high-pressure turbine of a gas turbine engine having a turbine casing in which, in addition to nozzle vanes, blades are provided which are arranged on a rotor and preferably have a shroud. The tips of the rotor blades or the shroud are surrounded by stator shroud segments suspended in the turbine casing, and a clearance is formed between the tips and stator shroud segments with a width which is controlled by a heating-cooling channel system, the channel system being supplied with a stream of air from the compressor section of the gas turbine and passing the combustion chamber of the gas turbine to reach, via a plurality of metering holes, an annulus bounded by a stator ring. For background art, reference is made to DE 30 40 594 C2. A clearance control system commonly includes a turbine casing in which, in addition to nozzle vanes, rotor blades are provided, with the rotor blades being arranged on a rotor and having a shroud surrounded by stator shroud segments. The stator shroud segments are suspended in the turbine casing and form a clearance with a width which is controlled by a heating-cooling channel system.
Gas turbines find widest use in aircraft gas turbine engines. Unlike in stationary applications, these engines are subjected to frequent load alternations and varying environmental constraints, causing dissimilar thermal expansions of the turbine casing on the one hand and of the turbine rotor on the other. The size of the clearance between the turbine blade tips and the surrounding turbine casing varies accordingly. For maximum gas turbine efficiency, it is desired to minimize the clearance size. In larger or more sophisticated gas turbine engines, active clearance control systems are used for this purpose, while in smaller engines, passive clearance control systems, like the simple one disclosed in DE 30 40 594 C2 cited above, may be sufficient.
In passive clearance control systems, especially in gas turbine engines with shrouded rotor blades, the turbine rotor clearance described above is determined, in cruise flight, by what is referred to as a hot reslam characteristic. As an engineer skilled in the art recognizes, the term "hot reslam" denotes hot re-acceleration of the gas turbine engine as briefly outlined below. It is assumed that, after a long cruise flight, both the turbine casing and the rotor section, i.e. the turbine rotor disk(s) bearing the rotor blades, are thoroughly heated to an elevated temperature level. When the engine reverts to low-load operation, e.g. idle speed, the turbine casing with its relatively thin wall section will cool faster than the rotor disks with their comparatively massive shapes. At some time in this process, therefore, the rotor disks will still be hot and expanded by their high temperature, while the turbine casing has already cooled and shrunk back. If, at this time, a high-load operation, such as a maximum operation at take-off, is resumed, then centrifugal force will cause the individual turbine rotor disk(s) to expand even further toward the turbine casing and the tips to come into undesirable contact with the casing.
The hot reslam characteristic worsens as the turbine casing cools more in idle operation and shrinks relative to the rotor. The faster cooling of the turbine casing as compared with the rotor causes the shroud on the rotor blades to grind a groove into the surrounding stator shroud segments of the turbine casing when the hot reslam occurs as described. The unwelcome turbine rotor clearance becomes larger as grinding becomes more intensive in cruise flight.
In a hot reslam condition, during a low-load operation of the aircraft gas turbine engine, the turbine casing should be heated rather than cooled to prevent excessive shrinking of the casing. This would keep the rotor blade shroud from grinding into the stator shroud segments as described above. The object underlying the present invention is to improve a passive clearance control system such that, especially during an idle operation, the turbine casing is heated rather than cooled.
This object is achieved by providing a ring-shaped orifice plate interacting with metering holes and closing the holes to varying degrees depending on the temperature of the incoming air flow to the holes. As a result, a fraction of the hot gases flowing across the nozzle vanes and rotor blades is allowed, at least at a low turbine load, to reach the annulus and preferably its aft or downstream region via a gap between the stator ring and the shroud segments. Further advantageous embodiments and developments are also contemplated.
This invention accordingly provides measures for affecting the temperature of the air flow passing into the heating-cooling channel system. This air flow is actually derived from differing sources. When the metering holes are exposed by the orifice plate, compressor air bypassing the combustion chamber of the gas turbine is allowed to reach the heating-cooling channel system and cool the turbine casing in the usual manner. The compressor air is relatively cool as compared with the hot gas stream of the gas turbine. If the metering holes are closed by the ring-shaped orifice plate, however, then a fraction of the hot gas stream reaches the heating-cooling channel system, causing the turbine casing to be cooled less or even to be heated. This ensures, especially in the hot reslam case in low-load operation, that the clearance between the tips of the rotor blades or of their shroud and the stator shroud segments is large enough to provide ample space to accommodate further expansion of the turbine rotor disk under centrifugal force, or the approach of the rotor blade tips toward the turbine casing when the engine is subsequently pushed into high-load operation, so that the undesirable grinding of the tips into the stator shroud is prevented.
The situation just described, in which a fraction of the hot gas stream reaches the heating-cooling channel system, occurs particularly at the idling point or some other low-load operating point of the gas turbine. Another operating situation, when compressor air bypassing the gas turbine combustion chamber is routed or ducted to the heating-cooling channel system, is, by contrast, a high-load operating range of the gas turbine. This high-load operating range occurs, for example, during cruise flight or take-off of the aircraft powered by the gas turbine engine.
Mixed states between these two extreme states, of course, can occur as well when a small amount of relatively cold air bypassing the gas turbine combustion chamber and a tiny fraction of the hot gas stream both reach the heating-cooling channel system or the annulus. Depending on the detailed design of a clearance control system according to the present invention, therefore, any desired heated or cooled state of the turbine casing can be achieved in order to obtain the desired clearance size between the rotor blade tips and the stator shroud segments surrounding them.
This clearance control system is advantageously passive in nature so that the ring-shaped orifice plate moves to expose or close the metering holes automatically in response to prevailing constraints, i.e. in response to the temperature of the compressor air flow bypassing the combustion chamber. To enable the orifice plate to serve this function, it could be made bimetallic, although a clearance control system in accordance with the invention will be especially simple and reliable when the orifice plate is made of material having a higher coefficient of thermal expansion than the stator ring. The metering holes will then be either closed or exposed simply by the different thermal expansions of the stator ring on the one hand and of the orifice plate on the other. With such an arrangement, best results are achieved when the ring-shaped orifice plate is arranged inside the annulus and has a relatively large circumference and, hence, a relatively large length by which to expand under heat, considering that it extends across the entire diameter of the (aircraft) gas turbine engine.